Turbine case reinforcement in a gas turbine jet engine

ABSTRACT

In one embodiment, a low or high pressure turbine case is machined on its outside surface to form circumferential notches. The notches may coincide with internal locations of seals for the blades, or with “hot spots” that have been identified, for example. A stiffener ring may be shrunk with an interference fit into each notch through inducing temperature differentials between the ring and the case. The radially compressive circumferential force exerted by each ring can inhibit the low or high pressure turbine case from expanding as much as it would otherwise. In some applications, a stiffener ring can improve blade tip clearance or counterbalance “hot spots”, stiffen the case, improve case cooling, or other benefits, depending upon the particular application. In one embodiment, notches may be avoided. In an alternate embodiment, C-rings, or multiple segmented rings, may be coupled together by hydraulic, electrical, or other means and actuated by a controller to exert adjustable radially compressive circumferential force. Other embodiments are described and claimed.

CROSS REFERENCE TO RELATED APPLICATIONS

This application claims the benefit of U.S. Provisional Application No.60/571,701, filed on May 17, 2004, titled “METHOD AND SYSTEM FORIMPROVED BLADE TIP CLEARANCE IN A GAS TURBINE JET ENGINE.”

A nonprovisional U.S. application entitled “METHOD AND SYSTEM FORIMPROVED BLADE TIP CLEARANCE IN A GAS TURBINE JET ENGINE” is being filedconcurrently by L. James Cardarella, John Usherwood and Andres DelCampo, wherein the contributions by John Usherwood and Andres Del Campohave been assigned to Carlton Forge Works, a California corporation.

BACKGROUND

Since the development of the gas turbine jet engine, blade tip clearancewithin the interior of the casing has been a challenging problem. Bladetip and inter-stage sealing have taken on a prominent role in enginedesign since the late 1960's. This is because the clearance between theblade tips and surrounding casing tends to vary due primarily to changesin thermal and mechanical loads on the rotating and stationarystructures. On today's largest land-based and aero turbine engines, thehigh pressure turbine case (“HPTC”) and low pressure turbine case(“LPTC”) have such large diameters that they are more susceptible toexpanding excessively and becoming out-of-round, exacerbating the bladetip clearance problem.

Reduced clearance in both the HPTC and the LPTC can provide dramaticreductions in specific fuel consumption (“SFC”), compressor stall marginand engine efficiency, as well as increased payload and mission rangecapabilities for aero engines. Improved clearance management candramatically improve engine service life for land-based engines andtime-on-wing (“TOW”) for aero engines. Deterioration of exhaust gastemperature (“EGT”) margin is the primary reason for aircraft engineremoval from service. The Federal Aviation Administration (“FAA”)certifies every aircraft engine with a certain EGT limit. EGT is used toindicate how well the HPTC is performing. Specifically, EGT is used toestimate the disk temperature within the HPTC. As components degrade andclearance between the blade tips and the seal on the interior of thecasing increase, the engine has to work harder (and therefore runshotter) to develop the same thrust. Once an engine reaches its EGTlimit, which is an indication that the high pressure turbine disk isreaching its upper temperature limit, the engine must be taken down formaintenance. Maintenance costs for major overhauls of today's largecommercial gas turbine jet engines can easily exceed one milliondollars.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a schematic diagram of the overall structure of a typicalgas turbine jet engine.

FIG. 2 shows a sectional schematic diagram of a low pressure turbinecase of a typical gas turbine jet engine.

FIG. 3 shows a sectional schematic diagram of the low pressure turbinecase of FIG. 2 fitted with stiffener rings in an embodiment of thepresent description.

FIG. 4 shows a sectional schematic diagram of Section A of the lowpressure turbine case of FIG. 3, showing the stiffener ring about to beseated in an embodiment of the present description.

FIG. 5 shows a sectional schematic diagram of a section of a lowpressure turbine case showing the stiffener ring about to be seated inanother embodiment of the present description.

FIG. 6 shows a sectional schematic diagram of a section of a lowpressure turbine case showing the stiffener ring seated in anotherembodiment of the present description.

FIG. 7 shows a sectional schematic diagram of a section of a lowpressure turbine case showing the stiffener ring seated in anotherembodiment of the present description.

FIG. 8 shows an improvement in clearance under load in an embodiment ofthe present description.

FIGS. 9A, 9B, and 9C show sectional schematic diagrams of a section of alow pressure turbine case having the stiffener ring positioned on thelow pressure turbine case with a hydraulic nut and secured with alocking nut in another embodiment of the present description.

FIG. 10 shows a schematic diagram of a low pressure turbine case havingstiffener rings actuated by hydraulic, electric, or other means inanother embodiment of the present description.

FIG. 11 shows a schematic cross-sectional diagram of a low pressureturbine case having stiffener rings.

DETAILED DESCRIPTION OF THE DRAWINGS

Referring now to the Figures, in which like reference numerals and namesrefer to structurally and/or functionally similar elements thereof, FIG.1 shows a schematic diagram of the overall structure of a typical gasturbine jet engine. Referring now to FIG. 1, Gas Turbine Jet Engine 100has Fan 102 for air intake within Fan Frame 104. High PressureCompressor Rotor 106 and its attached blades and stators force air intoCombustor 108, increasing the pressure and temperature of the inlet air.High Pressure Turbine Rotor 110 and its accompanying blades and statorsare housed within High Pressure Turbine Case 112. Low Pressure TurbineRotor 114 and its accompanying blades and stators are housed within LowPressure Turbine Case 116. The turbine extracts the energy from the highpressure, high-velocity gas flowing from Combustor 108 and istransferred to Low Pressure Turbine Shaft 118.

FIG. 2 shows a sectional schematic diagram of a low pressure turbinecase of a typical gas turbine jet engine. Referring now to FIG. 2,Centerline 202 runs through the center of Low Pressure Turbine Case 204(shown in cross-section). Rotor 206 (shown in cross-section) has Blade208 attached thereto and rotates on an axis of rotation along centerline202. One skilled in the art will recognize that many more blades andstators would normally be present within Low Pressure Turbine Case 204.Only one Blade 208 is shown for simplicity.

Labyrinth seal designs vary by application. Sometimes the labyrinthseals are located on the blade tips, and sometimes they are located onthe inside diameter of the cases as shown in FIG. 2. Labyrinth Seals 210(shown in cross-section) line the inside diameter of Low PressureTurbine Case 204 forming a shroud around each rotating Blade 208,limiting the air that spills over the tips of Blades 208. The shape ofLabyrinth Seals 210 is designed to create air turbulence between thetips of each Blade 208 and the corresponding Labyrinth Seal 210. The airturbulence acts as a barrier to retard air from escaping around the tipsof Blades 208. It is appreciated that seals performing similar functionsare often referred to by other names. Blade Tip Clearance 212, definedas the distance between the tip of Blade 208 and Labyrinth Seal 210,will vary over the operating points of the engine. The mechanisms behindBlade Tip Clearance 212 variations come from the displacement ordistortion of both static and rotating components of the engine due to anumber of loads on these components and expansion due to heat.Axis-symmetric clearance changes are due to uniform loading(centrifugal, thermal, internal pressure) on the stationary or rotatingstructures that create uniform radial displacement. Centrifugal andthermal loads are responsible for the largest radial variations in BladeTip Clearance 212.

Wear mechanisms for Labyrinth Seal 210 can be generally categorized intothree major categories: rubbing (blade incursion), thermal fatigue, anderosion. Engine build clearances in both high pressure and low pressureturbine cases are chosen to limit the amount of blade rubbing. Studieshave shown that improved blade tip clearances in the high pressure andlow pressure turbine cases can result in significant life cycle cost(“LCC”) reductions.

As a cold engine is started, a certain amount of Blade Tip Clearance 212exists between each Labyrinth Seal 210 and the tip of Blades 208. BladeTip Clearance 212 is rapidly diminished as the engine speed is increasedfor takeoff due to the centrifugal load on Rotor 206 as well as therapid heating of Blades 208, causing the rotating components to growradially outward. Meanwhile, Low Pressure Turbine Case 204 expands dueto heating but at a slower rate. This phenomenon can produce a minimumBlade Tip Clearance 212 “pinch point.” As Low Pressure Turbine Case 204expands due to heating after the pinch point, Blade Tip Clearance 212increases. Shortly after Low Pressure Turbine Case 204 expansion, Rotor206 begins to heat up (at a slower rate than Low Pressure Turbine Case204 due to its mass) and Blade Tip Clearance 212 narrows. As the engineapproaches the cruise condition, Low Pressure Turbine Case 204 and Rotor206 reach thermal equilibrium and Blade Tip Clearance 212 remainsrelatively constant.

There can be tremendous benefit in narrowing Blade Tip Clearance 212during the cruise condition. This is often where the greatest reductionin SFC can be gained (longest part of the flight profile). On the otherhand, rubbing is generally to be avoided. Minimal clearance typically ismaintained at takeoff to ensure thrust generation as well as keeping EGTbelow its established limit. Hence, it has been the goal of many controlsystems to attempt to maintain a minimal Blade Tip Clearance 212 whileavoiding rubbing over the entire flight profile.

Engine temperatures generally play a large role in determining theoperational Blade Tip Clearances 212. Gas turbine performance,efficiency, and life are directly influenced by Blade Tip Clearances212. Tighter Blade Tip Clearances 212 can reduce air leakage over thetips of Blades 208. This can increase turbine efficiency and permit theengine to meet performance and thrust goals with less fuel bum and lowerrotor inlet temperatures. Because the turbine runs at lowertemperatures, while producing the same work, hot section components canhave increased cycle life. The increased cycle life of hot sectioncomponents can increase engine service life (TOW) by increasing the timebetween overhauls.

Engine SFC and EGT are generally directly related to HPTC blade tipclearance. One study has shown that for every 0.001 inch increase inHPTC blade tip clearance, SFC increases approximately 0.1%, while EGTincreases one ° C. Therefore, it is believed that a 0.010 inch HPTCblade tip clearance decrease may roughly produce a one % decrease in SFCand a ten ° C. decrease in EGT. Military engines generally show slightlygreater HPTC blade tip clearance influence on SFC and EGT due to theirhigher operating speeds and temperatures over large commercial engines.Improvements of this magnitude may produce large savings in annual fueland engine maintenance costs amounting to over hundreds of millions ofdollars per year.

Reducing fuel consumption may also reduce aero engine total emissions.Recent estimates indicate that Americans alone now fly 764 million tripsper year (2.85 airline trips per person). The energy used by commercialaircraft has nearly doubled over the last three decades. The increasedfuel consumption accounts for thirteen % of the total transportationsector emissions of carbon dioxide (CO₂). Modem aero engine emissionsare made up of over seventy-one % CO₂ with about twenty-eight % water(H₂O) and 0.3% nitrogen oxide (NO₂) along with trace amounts of carbonmonoxide (CO), sulfur dioxide (SO₂), etc. Air transport accounts for2.5% (600 million tons) of the world's CO₂ Production. Emissions fromland-based engines, primarily for power generation, contributes amountsin addition to these totals. Clearly a reduction in fuel bum cansignificantly reduce aero and land-based engine emissions.

Current large commercial engines have cycle lives (defined as the timebetween overhauls) that vary significantly, ranging typically between3,000 to 10,000 cycles. The cycle life is primarily determined by howlong the engine retains a positive EGT margin. New engines or newlyoverhauled engines are shipped with a certain cold build blade tipclearance which increases with time. As the engine operating clearancesincrease, the engine generally works harder (hotter) to produce the samework and is therefore less efficient. This increase in operatingtemperature, particularly takeoff EGT, can further promote thedegradation of hot section components due to thermal fatigue. It isbelieved that retaining engine takeoff EGT margin by maintaining tightblade tip clearances can dramatically increase engine cycle life. Thiscould also lead to huge savings in engine maintenance over a period ofyears due to the large overhaul costs.

Previous attempts at blade tip clearance management can generally becategorized by two control schemes, active clearance control (“ACC”) andpassive clearance control (“PCC”). PCC is defined as any system thatsets the desired clearance at one operating point, namely the mostsevere transient condition (e.g., takeoff, re-burst, maneuver, etc.).ACC, on the other hand, is defined as any system that allows independentsetting of a desired blade tip clearance at more than one operatingpoint. The problem with PCC systems is that the minimum clearance, thepinch point, that the system must accommodate often leaves an undesiredlarger clearance during the much longer, steady state portion of theflight (i.e., cruise).

Typical PCC systems include better matching of rotor and stator growththroughout the flight profile, the use of abradables to limit blade tipwear, the use of stiffer materials and machining techniques to limit orcreate distortion of static components to maintain or improve shroudroundness at extreme conditions, and the like. Engine manufacturersbegan using thermal ACC systems in the late 1970's and early 1980's.These systems utilized fan air to cool the support flanges of the HPTC,reducing the case and shroud diameters, and hence blade tip clearance,during cruise conditions.

It is believed that all of the approaches described above havesignificant problems associated with them. Some are quite expensive,others achieve little results, especially during cruise where thegreatest advantages are gained, or require actuation through the casedue to the lack of current high temperature actuator capabilities, whichraise secondary sealing issues and added weight and mechanicalcomplexity.

FIG. 3 shows a sectional schematic diagram of the low pressure turbinecase of FIG. 2 fitted with stiffener rings in an embodiment of thepresent description. FIG. 11 shows a cross-sectional schematic diagramof the low pressure turbine case of FIG. 2 fitted with stiffener ringsin an embodiment of the present description. Referring now to FIGS. 3,11, one or more features of the present description may be applied toexisting gas turbine jet engines, or may be incorporated into the designand build of new gas turbine jet engines, for a variety of applicationsincluding aviation, marine and land-based engines. Features of thepresent description are applicable to the HPTC as well as the LPTC, andthe description and figures in relation to the LPTC also apply equallyto the HPTC and are not limited to the LPTC.

Notches 302, which may be of several different geometries as describedin detail below, are manufactured circumferentially, typically throughmachining, into the outside diameter of Low Pressure Turbine Case 204 tocoincide with one or more locations of the Labyrinth Seals 210. Inaddition to locations corresponding to one or more of the locations ofthe Labyrinth Seals 210, notches may be machined circumferentially inlocations corresponding to “hot spots” that have been identified in LowPressure Turbine Case 204 through computer modeling, through monitoringsurface temperatures, or through visual inspections for cracks when theengine is overhauled. For existing engines, Low Pressure Turbine Case204 is typically removed in order to repair cracks resulting from thethese “hot spots”. After such repairs, groves may then be appliedthrough a weld repair through machining. The external rings would thenbe shrink interference fit in the grooves. It is appreciated that thestiffener rings may be located at other positions of a turbine case,depending upon the particular application It is further appreciated thatsizes, dimensions, shapes, materials and clearances may vary, dependingupon the particular application.

In one embodiment, Stiffener Rings 304 (shown in cross section in FIG.3) are shrink interference fit into each Notch 302 so that the Stiffenerring 304 encircles the circumferential Notch 302 as shown in FIG. 11.Since Low Pressure Turbine Case 204 is conical in shape, each StiffenerRing 304 may have a different diameter. In each case, the insidediameter of each Stiffener Ring 304 may be slightly less than theoutside diameter of its corresponding Notch 302. Each Stiffener Ring 304is heated, starting with the largest diameter Stiffener Ring 304.Heating causes each Stiffener Ring 304 to expand, increasing the insidediameter to a diameter that is greater than the outside diameter of itscorresponding Notch 302. Once positioned in Notch 302, Stiffener Ring304 is allowed to cool, which shrinks with an interference fit into itscorresponding Notch 302.

FIG. 4 shows a sectional schematic diagram of Section A of the lowpressure turbine case of FIG. 3, showing the stiffener ring about to beseated in an embodiment of the present description. Referring now toFIG. 4, Notch 302 is manufactured circumferentially with a reverse taperrelative to the taper of the low pressure turbine case in oneembodiment. Angle 402 for the taper will vary from case to case, rangingfrom just greater than 0° for a cylindrical case to an appropriatedegree that would depend upon the specific geometry of a conical case.Stiffener Ring 304 may be machined circumferentially on its insidediameter to match this same taper. Even though Stiffener Ring 304 isshrink interference fit onto Low Pressure Turbine Case 204, the tapercan add extra security so that Stiffener Ring 304 is inhibited fromslipping axially on Low Pressure Turbine Case 204. If Notch 302 wasmanufactured flat without the taper, there may be an increasedpossibility of slippage in some applications. When Stiffener Ring 304has been heated it expands, giving rise to Ring Clearance 404, enablingStiffener Ring 304 to be positioned as shown against Heel 406 of Notch302. As Stiffener Ring 304 cools, it shrinks in diameter and seatsitself circumferentially into Notch 302. At ambient temperature, due tothe smaller diameter of the inner surface of the Stiffener Ring 304 tothe diameter of the outer surface of the Notch 302, a shrink with aninterference fit results, with radially compressive circumferentialforce being applied to Low Pressure Turbine Case 204 by Stiffener Ring304, and tensile circumferential force is applied to Stiffener Ring 304by Low Pressure Turbine Case 204. In one embodiment, the radiallycompressive forces may be centered on the axis of rotation defined bycenter line 202 as schematically shown by arrows in FIG. 11. In oneembodiment, the radially compressive forces are applied continuouslyaround the entire circumference of the Notch 302 and the Turbine Case204 without interruption.

In one example, Low Pressure Turbine Case 204 may be fifty inches inoutside diameter at the portion where Blade 208 and Labyrinth Seal 210are located. In one embodiment, the Stiffener Ring 304 may be fabricatedas a solid, unitary or one-piece, continuous or seamless member forgedor machined in a closed loop shape. In another embodiment, the StiffenerRing 304 may be fabricated using an open loop-shaped member and bondingthe ends together by welding, for example, to form a closed loop shape.Low Pressure Turbine Case 204 is made of nickel-based super alloy, suchas Inconel 718, as is Stiffener Ring 304 through a forging process.Super alloy Inconel 718 is a high-strength, complex alloy that resistshigh temperatures and severe mechanical stress while exhibiting highsurface stability, and is often used in gas turbine jet engines. It isappreciated that the stiffener ring and the turbine case may be made ofa variety of materials, depending upon the particular application.Heating Stiffener Ring 304 to a calculated temperature will causeStiffener Ring 304 to expand, yielding an appropriate Ring Clearance 404when Low Pressure Turbine Case 204 is at ambient air temperature ofapproximately seventy ° F. Alternatively, Low Pressure Turbine Case 204may be cooled with liquid nitrogen or other means to a calculatedtemperature to cause Low Pressure Turbine Case 204 to shrink indiameter, yielding an appropriate Ring Clearance 404 when Stiffener Ring304 is at ambient air temperature of approximately seventy ° F.Alternatively, an appropriate Ring Clearance 404 may be achieved througha combination of cooling Low Pressure Turbine Case 204 and heatingStiffener Ring 304, each to various calculated temperatures. Increasingor decreasing the inside diameter of Stiffener Ring 304 will result inmore or less radially compressive circumferential force and tensilestress as required for a particular application, and within the stresslimits of the material that Stiffener Ring 304 is made from.

In addition, the machining for Low Pressure Turbine Case 204 may be donein a first direction, such as radially, and the machining for StiffenerRing 304 may be done in a second direction, such as axially, which ismore or less perpendicular to the first direction. Since machiningleaves a spiral, or record, continuous groove on the machined surfaces,the grooves on each surface will align in a cross-hatch manner to eachother, increasing the frictional forces between the two surfaces andreducing the potential for movement of Stiffener Ring 304 within Notch302, including axial or rotational movement. The plurality of grooves onStiffener Ring 304, which may be made of a nickel-base super alloy forexample, may be harder than the plurality of grooves on Notch 302 of LowPressure Turbine Case 204, which is typically made of titanium, or inother low pressure turbine casings, possibly steel or aluminum. Thenickel-base super alloy grooves can dent into or form an indentation inthe softer titanium, steel, or aluminum grooves. Alternatively,Stiffener Ring 304 may simply be spot welded in one or more locations toNotch 302, or bolted to one or more flanges secured to Notch 302, tokeep Stiffener Ring 304 from spinning or otherwise moving in relation toNotch 302. Machining in cross directions may not be needed in this case.

By thus positioning Stiffener Rings 304 in the manner described, BladeTip Clearance 212 may be improved in some applications, especiallyduring cruise operation of the engine in some applications. An enginedesigner may as a result, design the engine to have a reduced blade tipclearance than may otherwise be appropriate for a given engine designabsent such stiffener rings. It is also appreciated that other ordifferent benefits, advantages, improvements or other features may beutilized alone or in combination, depending upon the particularapplication. In one application, the radially compressivecircumferential force (represented by arrows in FIG. 11) applied by theStiffener Rings 304 can prevent Low Pressure Turbine Case 204 fromexpanding due to heat as much as it would otherwise expand. In oneaspect, the Stiffener Rings 304 function as a girdle for the TurbineCase 204, to inhibit expansion or going out of round and otherwisereinforce the Turbine Case 204. Stiffener Rings 304 may be made of thesame material as Low Pressure Turbine Case 204, or may be made of adifferent material with a lower coefficient of thermal expansion, whichwould increase the radially compressive circumferential force appliedover that of a stiffener ring of the same material as the case as thetemperature rises. The compressive forces may be sufficient to form anindentation in the turbine case such as in the Notch 302.

In many engine designs, heat is mainly dissipated from the outsidesurface area of Low Pressure Turbine Case 204 by convection. Anotherbenefit which may be achieved by adding Stiffener Rings 304 to LowPressure Turbine Case 204 is that heat may be dissipated at a greaterrate because Stiffener Rings 304 can act as cooling fins, which canresult in cooler operating temperatures within Low Pressure Turbine Case204. This cooling may also contribute to less expansion and smallerBlade Tip Clearance 212. Also, Stiffener Rings 304 can help to maintainroundness of Low Pressure Turbine Case 204. Again, it is appreciatedthat other or different benefits, advantages, improvements or otherfeatures may be utilized alone or in combination, depending upon theparticular application.

FIG. 5 shows a sectional schematic diagram of a section of a lowpressure turbine case showing the stiffener ring about to be seated inanother embodiment of present description. Referring now to FIG. 5,Notch 502 is machined circumferentially with a chevron shape in oneembodiment. Angle 508 may vary by application. Stiffener Ring 504 ismachined circumferentially on its inside diameter to match this samechevron shape. Even though Stiffener Ring 504 is shrink interference fitonto Low Pressure Turbine Case 204, the chevron shape can add extrasecurity to inhibit the Stiffener Ring 304 from slipping off of LowPressure Turbine Case 204. When Stiffener Ring 504 has been heated itexpands, giving rise to Ring Clearance 404, enabling Stiffener Ring 504to be positioned as shown against Heel 506 of Notch 502. As StiffenerRing 504 cools, it shrinks in diameter and seats itselfcircumferentially into Notch 502. At ambient temperature, due to thesmaller inside diameter of Stiffener Ring 504 to the outside diameter ofNotch 502, a shrink with an interference fit results, with radiallycompressive circumferential force being applied to Low Pressure TurbineCase 204 by Stiffener Ring 504, and tensile circumferential force isapplied to Stiffener Ring 504 by Low Pressure Turbine Case 204.

FIG. 6 shows a sectional schematic diagram of a section of a lowpressure turbine case showing the stiffener ring seated in anotherembodiment of the present description. Referring now to FIG. 6, for aeroapplications, where added weight to the engine is a concern, StiffenerRing 604 is manufactured to have a profile that, when seated as shown inFIG. 6, is substantially flush with the outer surface of Low PressureTurbine Case 204. Notch 302 with a reverse taper as shown in FIG. 4 ismachined into Low Pressure Turbine Case 204. In addition, based on theengine to be designed or to be retrofitted, Notch 302 may be machineddeeper, and/or wider, and Stiffener Ring 604 given added depth, and/orwidth, in order to meet the radially compressive and tensilecircumferential stress requirements.

FIG. 7 shows a sectional schematic diagram of a section of a lowpressure turbine case showing the stiffener ring seated in anotherembodiment of the present description. Referring now to FIG. 7, for aeroapplications, where added weight to the engine is a concern, StiffenerRing 704 is manufactured to have a profile that, when seated as shown inFIG. 6, is substantially flush with the outer surface of Low PressureTurbine Case 204. Notch 502 with a chevron shape as shown in FIG. 5 ismachined into Low Pressure Turbine Case 204. In addition, based on theengine to be designed or to be retrofitted, Notch 502 may be machineddeeper and/or wider, and Stiffener Ring 704 given added depth, and/orwidth, in order to meet the radially compressive and tensile stressrequirements. In addition to aero or aviation applications, it isappreciated that flush embodiments as well as other embodiments may beutilized in land-based and marine applications as well.

One skilled in the art will recognize that, in addition to the reversetaper and chevron designs for the notch and stiffener ring as shown inFIGS. 4-7, various other designs may be utilized to accomplish the sameor similar or different goals. For example, the notch may have one ormore ridges and channels, angular or undulating, that will match up withone or more channels and ridges, angular or undulating, on the insidesurface of the stiffener ring. Alternatively, the notch and stiffenerring may have an inverted chevron shape. In other embodiments, a notchmay not be utilized. Many other such shapes may be envisioned withoutdeparting from the scope of the present description.

FIG. 8 shows the improvement in blade tip clearance under load in anembodiment of the present description. Referring now to FIG. 8,Stiffener Ring 304 as shown in FIG. 4 has been shrink interference fitonto Low Pressure Turbine Case 204, and the engine is now under load,such as during cruise operation. Labyrinth Seal 210 and Low PressureTurbine Case 204 with Inner Surface 802 and Outer Surface 804 aredepicted with solid lines in the positions they would be in withoutStiffener Ring 304. Low Pressure Turbine Case 204 would have expanded indiameter, and Labyrinth Seal 210 would have moved away from Blade 208,giving rise to a wider Blade Tip Clearance 212. However, due to theradially compressive force exerted by Stiffener Ring 304 on Low PressureTurbine Case 204, Labyrinth Seal 210 is in the position indicated inphantom as 210′, and Ring 304, Inner Surface 802 and Outer Surface 804of Low Pressure Turbine Case 204 are in the positions indicated inphantom as 304′, 802′, and 804′, thus reducing Blade Tip Clearance 212′.

Thus, in one aspect of the present description, the amount of expansionthat would normally occur due to heating in the LPTC and the HPTC, isreduced, and consequently blade tip clearance may be improved. As statedabove, increased blade tip clearance can accelerate the effects of lowcycle fatigue and erosion due to increased temperatures in the HPTC andLPTC, and degrade EGT margin and engine life. In general, for large gasturbine engines, it is believed that blade tip clearance reductions onthe order of 0.010 inch can produce decreases in SFC of one % and EGT often ° C. It is believed that improved blade tip clearance of thismagnitude can produce fuel and maintenance savings of over hundreds ofmillions of dollars per year. Reduced fuel bum can also reduce aircraftemissions, which currently account for thirteen % of the total U.S.transportation sector emissions of CO₂. In another aspect, blade tipclearances can be reduced at cruise condition to make a significantimpact on SFC and EGT margin and improve turbine efficiency. Moreover,the increased outer surface area of the HPTC and LPTC due to thestiffener rings can, in certain embodiments, increase cooling and resultin lower internal temperatures which can lengthen the cycle life of theengine. In yet another aspect, an increase in payload per engine may beachieved due to the improvement in blade tip clearance. Additionalpounds of freight may be transported per takeoff and landing. It isfurther appreciated that features of the present description couldreadily replace expensive passive clearance control options. It isappreciated that reductions in one or more of out-of-roundness, bladetip clearance, SFC, EGT or polluting emissions may be achieved utilizingone or more features herein described. For example, fabricating astiffener ring from a material having a lower coefficient of thermalexpansion than that of the turbine case material, may facilitateachieving one or more of these or other reductions. Similarly, it isappreciated that one or more of these reductions or other benefits maybe achieved fabricating a turbine case and stiffener ring of the samematerial.

FIGS. 9A, 9B, and 9C show sectional schematic diagrams of a section of alow pressure turbine case having the stiffener ring positioned on thelow pressure turbine case with a hydraulic nut and secured with alocking nut in another embodiment. Referring now to FIG. 9A, StiffenerRing 904 is sized to fit without pressure in a location near an internalBlade 208 and Labyrinth Seal 210, or previously identified “hot spot”,and placed in position there. Next, a Hydraulic Nut 902 is threadablymounted to Low Pressure Turbine Case 204. Hydraulic Nut 902 has Piston906 which engages with Stiffener Ring 904.

In FIG. 9B, Piston 906 has extended from Hydraulic Nut 902, pushingStiffener Ring 904 toward the larger diameter end of Low PressureTurbine Case 204, thus positioning Stiffener Ring 904 in the optimumlocation in relation to the internal Blade 208 and Labyrinth Seal 210and resulting in an interference fit. The amount that Piston 906 isextended by Hydraulic Nut 902 is calculated to produce a desiredcompressive circumferential force by Stiffener Ring 904.

In FIG. 9C, Hydraulic Nut 902 has been removed, and Locking Nut 908 hasbeen threadably attached in its place onto Low Pressure Turbine Case204. Retainer 910 of Locking Nut 908 engages with Stiffener Ring 904,thus securing Stiffener Ring 904 in place. This process is repeated foras many stages as required based upon turbine design. This embodimentmay add excessive weight and would most likely be best suited for landbased applications where weight is not of such concern.

FIG. 10 shows a schematic diagram of a low pressure turbine case havingstiffener rings actuated by hydraulic, electric, or other means inanother embodiment of the present description. Referring now to FIG. 10,Low Pressure Turbine Case 1000 has Stiffener C-Rings 1004 positioned atpredetermined locations to coincide with blade/labyrinth seals and/or“hot spots”. In this embodiment, Stiffener C-Rings 1004 are not shrinkinterference fit onto Low Pressure Turbine Case 1000. A notch for eachStiffener C-Ring 1004 may still be machined into Low Pressure TurbineCase 1000, but the stiffener rings are c-rings rather than continuousrings. Each end of Stiffener C-Ring 1004 is linked to an Actuator Means1002, which when actuated, pulls each end of Stiffener C-Ring 1004together, exerting compressive force including radially compressiveforce on Low Pressure Turbine Case 1000. The inside surface of eachStiffener C-Ring 1004, or the notch surface, or both, may be coated withTeflon® or some other lubricating substance to facilitate slippage whentightened.

Each Actuator Means 1002 is connected to Controller 1008 throughElectrical/Electronic Connections 1006. Controller 1008 receivestemperature readings from multiple temperature sensors located near eachStiffener C-Ring 1004 (not shown). It is also possible to derive theLPTC temperature from EGT temperature readings and use these readingsfor feedback to Controllers 1008. As the temperatures being monitoredthroughout Low Pressure Turbine Case 1000 rise, Controller 1008processes the temperature data and determines how much each of the endsof each Stiffener C-Ring 1004 need to be pulled together by eachActuator Means 1002 in order to exert the proper compressivecircumferential force on Low Pressure Turbine Case 1000 to provide asuitable benefit such as maintaining an optimum blade tip clearance orcounterbalancing a “hot spot”, for example.

In an alternate embodiment, instead of a c-ring, a chain-like multiplesegmented ring may be coupled together by Actuator Means 1002. Inanother embodiment, the stiffener rings may be made of a strip ofnon-metallic material, such as Kevlar®. The inside surface of theKevlar®, or the notch surface, or both may also be coated with Teflon®or some other lubricating substance to facilitate slippage whentightened.

Having described various features, it will be understood by thoseskilled in the art that many and widely differing embodiments andapplications will suggest themselves without departing from the scope ofthe present description.

1. A method, comprising: encircling an outer circumferential surface ofa turbine case of a gas turbine jet engine using an innercircumferential surface of a stiffener ring; and applying radiallycompressive forces to said outer circumferential surface of said turbinecase, along the length of the circumference of said innercircumferential surface, using said stiffener ring encircling saidturbine case.
 2. The method of claim 1 wherein said radially compressiveforce applying includes shrink interference fitting said stiffener ringinner circumferential surface to said outer circumferential surface ofsaid turbine case.
 3. The method of claim 1 wherein said radiallycompressive force applying includes seating said stiffener ring innercircumferential surface within a notch defined by said outercircumferential surface of said turbine case and shaped to secure saidstiffener ring against displacement in a direction longitudinal to saidturbine case.
 4. The method of claim 1 wherein said turbine casesurrounds a turbine adapted for rotation within said turbine case alongan axis of rotation wherein said radially compressive forces aredirected to a center located on said axis of rotation.
 5. The method ofclaim 4 wherein said turbine case has a seal encircling tips of saidturbine blades of said turbine and wherein said radially compressiveforce applying confines the clearance between said seal and said bladetips to be within a predetermined range.
 6. The method of claim 5wherein said turbine case is formed of a first material and wherein saidstiffener ring is formed of a second material that is different fromsaid first material of said turbine case, said second material having alower coefficient of thermal expansion than said first material of saidturbine.
 7. The method of claim 5 wherein said radially compressiveforces applied to said outer circumferential surface of said turbinecase, form an indentation in said outer circumferential surface alongthe length of the circumference of said inner circumferential surface ofthe stiffener ring, as the temperature of the turbine case rises.
 8. Themethod of claim 1 further comprising redesigning the engine to reduceblade tip clearance as compared to the blade tip clearance of saidengine absent said radially compressive force applying.
 9. The method ofclaim 1 wherein said applying radially compressive forces permits atleast one of the following to be reduced during operation of saidengine: a) turbine case out-of-roundness; b) specific fuel consumption;c) clearance between an inner surface of said turbine case and bladetips of said turbine; d) exhaust gas temperature; e) exhaust gaspollution.
 10. The method of claim 1 further comprising: encircling asecond outer circumferential surface of said turbine case of said gasturbine jet engine using a second inner circumferential surface of asecond stiffener ring; and applying radially compressive forces to saidsecond outer circumferential surface of said turbine case, along thelength of the circumference of said second inner circumferentialsurface, using said second stiffener ring encircling said turbine case.11. A method of operating a gas turbine jet engine, comprising: rotatinga turbine within a turbine case along an axis of rotation; and applyingradially compressive forces to an outer circumferential surface of saidturbine case using a stiffener ring encircling said turbine case, saidradially compressive forces being applied along the length of thecircumference of an inner circumferential surface of said stiffener ringand directed to a center positioned on said axis of rotation.
 12. Themethod of claim 11 wherein said turbine case has a seal encircling tipsof said turbine blades of said turbine and wherein said radiallycompressive force applying confines the clearance between said seal andsaid blade tips to be within a predetermined range as said turbinerotates within said turbine case.
 13. The method of claim 11 furthercomprising dissipating heat from said turbine case using said stiffenerring.
 14. The method of claim 11 further comprising redesigning theengine to reduce blade tip clearance as compared to the blade tipclearance of said engine absent said radially compressive forceapplying.
 15. The method of claim 11 wherein said applying radiallycompressive forces permits at least one of the following to be reducedduring operation of said engine: a) turbine case out-of-roundness; b)specific fuel consumption; c) clearance between an inner surface of saidturbine case and blade tips of said turbine; d) exhaust gas temperature;e) exhaust gas pollution.
 16. The method of claim 15 wherein saidturbine case is formed of a first material and wherein said stiffenerring is formed of a second material that is different from said firstmaterial of said turbine case, said second material having a lowercoefficient of thermal expansion than said first material of saidturbine.
 17. The method of claim 11 further comprising applying radiallycompressive forces to a second outer circumferential surface of saidturbine case using a second stiffener ring encircling said turbine case,said radially compressive forces being applied along the length of thecircumference of a second inner circumferential surface of said secondstiffener ring and directed to a center positioned on said axis ofrotation.
 18. A gas turbine jet engine, comprising: a turbine casehaving an outer circumferential surface; a turbine adapted to rotatealong an axis of rotation within said turbine case; and a stiffener ringhaving an inner circumferential surface adapted to apply radiallycompressive forces to said outer circumferential surface of said turbinecase, along the length of the circumference of said innercircumferential surface.
 19. The engine of claim 18 wherein saidstiffener ring is affixed to said turbine case with a shrinkinterference fitting which causes said stiffener ring to apply saidradially compressive forces to said outer circumferential surface ofsaid turbine case.
 20. The engine of claim 18 wherein said outercircumferential surface of said turbine case defines a notch adapted toreceive said stiffener ring and secure said stiffener ring againstdisplacement in a direction longitudinal to said turbine case.
 21. Theengine of claim 18 wherein said radially compressive forces are directedto a center located on said axis of rotation.
 22. The engine of claim 21wherein said turbine has turbine blade, each of which has a tip at adistal end of each blade, and wherein said turbine case has an innercircumferential surface which has a seal encircling said tips of saidturbine blades of said turbine and wherein said radially compressiveforce applied by said stiffener ring confines the clearance between saidseal and said blade tips to be within a predetermined range.
 23. Theengine of claim 21 wherein said outer circumferential surface of saidturbine case defines a notch adapted to receive said stiffener ring andsecure said stiffener ring against displacement in a directionlongitudinal to said turbine cases and wherein said notch is at alongitudinal location coinciding with said seal on said inner surface ofsaid turbine case.
 24. The engine of claim 18 wherein said turbine caseis formed of a first material and wherein said stiffener ring is formedof a second material that is different from said first material of saidturbine case, said second material having a lower coefficient ofexpansion than said first material of said turbine case.
 25. The engineof claim 22 further wherein said stiffener ring permits redesigning theengine to reduce blade tip clearance as compared to the blade tipclearance of said engine absent said radially compressive forceapplying.
 26. The engine of claim 22 wherein said stiffener ring permitsat least one of the following to be reduced during operation of saidengine: a) turbine case out-of-roundness; b) specific fuel consumption;c) clearance between an inner surface of said turbine case and bladetips of said turbine; d) exhaust gas temperature; e) exhaust gaspollution.
 27. The engine of claim 18 wherein said turbine case hassecond outer circumferential surface, and wherein said engine furthercomprises a second stiffener ring having a second inner circumferentialsurface adapted to apply radially compressive forces to a second outercircumferential surface of said turbine case, along the length of thecircumference of said second inner circumferential surface.
 28. Amethod, comprising: (a) machining at least one notch circumferentiallyat a predetermined location into an outer surface of a turbine case of agas turbine jet engine; and (b) seating a stiffener ring in each said atleast one notch through a shrink interference fit; wherein saidstiffener ring applies compressive circumferential force to said turbinecase.
 29. The method according to claim 28 wherein said seating furthercomprises: heating said stiffener ring to cause a first inside diameterof said stiffener ring to increase to a second inside diameter that islarger than an outside diameter of said at least one notch at an ambienttemperature; positioning said stiffener ring in said at least one notch;and allowing said stiffener ring to cool to said ambient temperature,causing said stiffener ring to decrease from said second inside diametertoward said first inside diameter, but resisted by said outside diameterof said at least one notch, giving rise to said shrink interference fit.30. The method according to claim 28 wherein said seating furthercomprises: cooling said turbine case to cause a first outside diameterof said at least one notch to decrease to a second outside diameter thatis smaller than an inside diameter of said stiffener ring at an ambienttemperature; positioning said stiffener ring in said at least one notch;and allowing said turbine case to heat up to said ambient temperature,causing said at least one notch to increase from said second outsidediameter toward said first outside diameter, but resisted by said insidediameter of said stiffener ring, giving rise to said shrink interferencefit.
 31. The method according to claim 28 wherein said seating furthercomprises: heating said stiffener ring to cause a first inside diameterof said stiffener ring to increase to a second inside diameter; coolingsaid turbine case to cause a first outside diameter of said at least onenotch to decrease to a second outside diameter that is smaller than saidsecond diameter of said stiffener ring; positioning said stiffener ringin said at least one notch; allowing said stiffener ring to cool to saidambient temperature; and allowing said turbine case to heat up to saidambient temperature; wherein said stiffener ring decreases from saidsecond inside diameter toward said first inside diameter, and said atleast one notch increases from said second outside diameter toward saidfirst outside diameter, giving rise to said shrink interference fit. 32.The method according to claim 28 wherein said machining furthercomprises: machining said at least one notch circumferentially into anouter surface of said turbine case at a location coinciding with alabyrinth seal on an inner surface of said turbine case.
 33. The methodaccording to claim 28 wherein said machining further comprises:machining said at least one notch circumferentially into an outersurface of said turbine case at a location coinciding with a hot spot ofsaid turbine case.
 34. The method according to claim 28 furthercomprising: machining said stiffener ring to a predetermined shape tomatch with a shape of said at least one notch.
 35. The method accordingto claim 34 wherein said notch machining comprises machining said atleast one notch circumferentially at said predetermined location intosaid outer surface of said turbine case with a reverse taper; andwherein said stiffener ring machining comprises machining said stiffenerring on an inside diameter to match said reverse taper of said at leastone notch.
 36. The method according to claim 34 wherein said notchmachining comprises machining said at least one notch circumferentiallyat said predetermined location into said outer surface of said turbinecase with a chevron shape; and wherein said stiffener ring machiningcomprises machining said stiffener ring on an inside diameter to matchsaid chevron shape of said at least one notch.
 37. The method accordingto claim 34 wherein said stiffener ring machining comprises machining atop surface of said stiffener ring so that when said stiffener ring isseated in said at least one notch, said top surface of said stiffenerring is flush with said outer surface of said turbine case.
 38. Themethod according to claim 34 wherein said stiffener ring machiningcomprises machining said stiffener ring from a nickel-base super alloy.39. The method according to claim 34 wherein said stiffener ringmachining comprises machining said stiffener ring from a material thatis different from a material of said turbine case, said material of saidstiffener ring having a lower coefficient of expansion than saidmaterial of said turbine case.
 40. A method according to claim 34wherein said notch machining comprises machining said at least one notchinto said outer surface of said turbine case in a first direction,wherein a plurality of grooves are formed and aligned on said outersurface in said first direction; and wherein said stiffener ringmachining comprises machining an inner surface of said stiffener ring ina second direction, wherein a plurality of grooves are formed andaligned on said inner surface in said second direction; wherein whensaid outer surface of said at least one notch and said inner surface ofsaid stiffener ring are seated together, said plurality of grooves onsaid outer surface of said at least one notch and said plurality ofgrooves on said inner surface of said stiffener ring align in across-hatch manner to each other, increasing the frictional forcesbetween said at least one notch and said stiffener ring and reducing thepotential for spinning of said stiffener ring within said at least onenotch.
 41. An apparatus for use in a gas turbine jet engine, theapparatus comprising: a turbine case having an outer surface whichdefines at least one notch machined circumferentially into said outersurface of said turbine case of said gas turbine jet engine at apredetermined location; and a stiffener ring seated in each said atleast one notch through a shrink interference fit; wherein saidstiffener ring applies compressive circumferential force to said turbinecase.
 42. The apparatus according to claim 41 further comprising: ameans for heating said stiffener ring to cause a first inside diameterof said stiffener ring to increase to a second inside diameter that islarger than an outside diameter of said at least one notch at an ambienttemperature, wherein after said stiffener ring is positioned in said atleast one notch, said stiffener ring is allowed to cool to said ambienttemperature, causing said stiffener ring to decrease from said secondinside diameter toward said first inside diameter, but resisted by saidoutside diameter of said at least one notch, giving rise to said shrinkinterference fit.
 43. The apparatus according to claim 41 furthercomprising: a means for cooling said turbine case to cause a firstoutside diameter of said at least one notch to decrease to a secondoutside diameter that is smaller than an inside diameter of saidstiffener ring at an ambient temperature, wherein after said stiffenerring is positioned in said at least one notch, said turbine case isallowed to heat up to said ambient temperature, causing said at leastone notch to increase from said second outside diameter toward saidfirst outside diameter, but resisted by said inside diameter of saidstiffener ring, giving rise to said shrink interference fit.
 44. Theapparatus according to claim 41 further comprising: a means for heatingsaid stiffener ring to cause a first inside diameter of said stiffenerring to increase to a second inside diameter; and a means for coolingsaid turbine case to cause a first outside diameter of said at least onenotch to decrease to a second outside diameter that is smaller than saidsecond diameter of said stiffener ring, wherein after said stiffenerring is positioned in said at least one notch, said stiffener ring isallowed to cool to said ambient temperature and said turbine case isallowed to heat up to said ambient temperature, causing said stiffenerring to decrease from said second inside diameter toward said firstinside diameter, and said at least one notch to increase from saidsecond outside diameter toward said first outside diameter, giving riseto said shrink interference fit.
 45. The apparatus according to claim 41wherein said predetermined location for machining said at least onenotch circumferentially into said outer surface of said turbine case isat a location coinciding with a labyrinth seal on an inner surface ofsaid turbine case.
 46. The apparatus according to claim 41 wherein saidpredetermined location for machining said at least one notchcircumferentially into said outer surface of said turbine case is at alocation coinciding with a hot spot of said turbine case.
 47. Theapparatus according to claim 41 wherein said stiffener ring furthercomprises a predetermined shape machined to match with a shape of saidat least one notch.
 48. The apparatus according to claim 41 wherein saidnotch has a reverse taper machined into said at least one notch; andwherein said stiffener ring has a matching reverse taper machined on aninside diameter of said stiffener ring.
 49. The apparatus according toclaim 41 further comprising: a chevron shape machined into said at leastone notch; and a matching chevron shape machined on an inside diameterof said stiffener ring.
 50. The apparatus according to claim 41 whereinsaid stiffener ring further comprises: a top surface of said stiffenerring machined so that when said stiffener ring is seated in said atleast one notch, said top surface of said stiffener ring is flush withsaid outer surface of said turbine case.
 51. The apparatus according toclaim 41 wherein said stiffener ring is machined from a nickel-basesuper alloy.
 52. The apparatus according to claim 41 wherein saidstiffener ring is machined from a material that is different from amaterial of said turbine case, said material having a lower coefficientof thermal expansion than said material of said turbine case.
 53. Theapparatus according to claim 41 further wherein said at least one notchof said turbine case has a machined outer surface which defines aplurality of grooves aligned in a first direction on said machined outersurface of said at least one notch; and wherein said stiffener ring hasa machined inner surface which defines a plurality of grooves aligned ina second direction on said machined inner surface of said stiffenerring; wherein when said outer surface of said at least one notch andsaid inner surface of said stiffener ring are interference shrink fittogether, said plurality of grooves on said outer surface of said atleast one notch and said plurality of grooves on said inner surface ofsaid stiffener ring align in a cross-hatch manner to each other,increasing the frictional forces between said at least one notch andsaid stiffener ring and reducing the potential for spinning of saidstiffener ring within said at least one notch.
 54. A method comprising:(a) machining at least one notch circumferentially at a predeterminedlocation into an outer surface of a turbine case of a gas turbine jetengine; (b) seating a stiffener ring in each said at least one notch,said stiffener ring having a first end and a second end; (c) linkingsaid first end and said second end of said stiffener ring to anactuator; and (d) actuating said actuator to pull said first and secondends of said stiffener ring together; wherein said stiffener ringapplies compressive circumferential force to said turbine case.
 55. Themethod according to claim 54 wherein said machining further comprisesmachining said at least one notch circumferentially into an outersurface of said turbine case at a location coinciding with a labyrinthseal on an inner surface of said turbine case.
 56. The method accordingto claim 54 wherein said machining further comprises machining said atleast one notch circumferentially into an outer surface of said turbinecase at a location coinciding with a hot spot of said turbine case. 57.The method according to claim 54 further comprising machining saidstiffener ring to a predetermined shape to match with a shape of said atleast one notch.
 58. The method according to claim 54 further comprisingmachining said stiffener ring from a nickel-base super alloy.
 59. Themethod according to claim 54 further comprising machining said stiffenerring from a material that is different from a material of said turbinecase, said material of said stiffener ring having a lower coefficient ofthermal expansion than said material of said turbine case.
 60. Themethod according to claim 54 further comprising: connecting a controllerto said actuator through an electrical connection; receiving in saidcontroller a plurality of temperature readings from a plurality oftemperature sensors located near said stiffener ring; and processing bysaid controller said plurality of temperature readings to determine howmuch to pull said first and second ends of said stiffener ring togetherby said actuator to exert a predetermined compressive circumferentialforce on said turbine case.
 61. The method according to claim 54 whereinsaid stiffener ring is a one of a c-ring, a chain like multiplesegmented ring, and a strip of non-metallic material.
 62. An apparatusfor use in a gas turbine jet engine, the apparatus comprising: a turbinecase having an outer surface which defines at least one notch machinedcircumferentially into said outer surface of said turbine case of thegas turbine jet engine at a predetermined location; a stiffener ringseated in each said at least one notch, said stiffener ring having afirst end and a second end; and an actuator, wherein said first andsecond ends are linked to said actuator and said actuator when actuatedis adapted to pull said first and second ends together; wherein saidstiffener ring applies compressive circumferential force to said turbinecase.
 63. The apparatus according to claim 62 wherein said turbine casehas an inner surface and a labyrinth seal on said inner surface andwherein said predetermined location for machining said at least onenotch circumferentially into said outer surface of said turbine case isat a location coinciding with said labyrinth seal on said inner surfaceof said turbine case.
 64. The apparatus according to claim 62 whereinsaid turbine case has a hot spot and wherein said predetermined locationfor machining said at least one notch circumferentially into said outersurface of said turbine case is at a location coinciding with said hotspot of said turbine case.
 65. The apparatus according to claim 62wherein said stiffener ring further comprises a predetermined shape tomatch with a shape of said at least one notch.
 66. The apparatusaccording to claim 62 wherein said stiffener ring is machined from anickel-base super alloy.
 67. The apparatus according to claim 62 whereinsaid stiffener ring is machined from a material that is different from amaterial of said turbine case, said material of said stiffener ringhaving a lower coefficient of thermal expansion than said material ofsaid turbine case.
 68. The apparatus according to claim 62 furthercomprising: a controller connected to said actuator through anelectrical connection; and a plurality of temperature sensors locatednear said stiffener ring, wherein said controller is adapted to receivea plurality of temperature readings from said plurality of temperaturesensors; wherein said controller is adapted to process said plurality oftemperature readings to determine how much to pull said first and secondends of said stiffener ring together by said actuator to exert apredetermined compressive circumferential force on said turbine case.69. The apparatus according to claim 62 wherein said stiffener ring is aone of a c-ring, a chain like multiple segmented ring, and a strip ofnon-metallic material.